This invention relates generally to substrates used in a high-temperature environment. More particularly, it relates to methods for protecting the substrates from damage in such an environment.
Various types of materials, such as metals and ceramics, are used for components which may be exposed to a high-temperature environment. Aircraft engine parts represent examples of these types of components. A variety of approaches have been used to raise the operating temperature at which the metal components can be used. For example, one approach involves the use of protective coatings on various surfaces of the component, e.g., a turbine engine airfoil. The coatings are usually ceramic-based, and are sometimes referred to as thermal barrier coatings or "TBC's".
The TBC's are typically used in conjunction with internal cooling channels within the airfoil, through which cool air is forced during engine operation. As an example, a pattern of cooling holes may extend from a relatively cool surface of an airfoil to a "hot" surface which is exposed to gas flow at combustion temperatures of at least about 1000.degree. C. The technique is sometimes referred to as "discrete hole film cooling". Cooling air, usually bled off from the engine's compressor, is typically bypassed around the turbine engine's combustion zone and fed through the cooling holes to the hot surface. The ratio of the cooling air mass flux (the product of air velocity times density) to the mass flux of the hot gas flowing along the hot surface (e.g., a combustion product) is sometimes referred to as the "blowing ratio". The cooling air forms a protective "film" between the hot surface and the hot gas flow, preventing melting or other degradation of the component, as described in U.S. Pat. No. 5,458,461 (C. P. Lee et al).
Film cooling performance may be characterized in several ways. One relevant indication of performance is known as the adiabatic wall film cooling effectiveness, sometimes referred to herein as the "cooling effectiveness". This particular parameter is related to the concentration of film cooling fluid at the surface being cooled. In general, the greater the cooling effectiveness, the more efficiently can the surface be cooled. A decrease in cooling effectiveness requires greater amounts of cooling air to try to maintain a certain cooling capacity. This requirement in turn diverts air away from the combustion zone. This can lead to other problems, such as greater air pollution resulting from non-ideal combustion, and less efficient engine operation. In the case of a turbine engine airfoil, effective film cooling requires that the film adhere to the hot surface of the airfoil, with as little mixing as possible with the hotter combustion gases.
A method for increasing cooling effectiveness is described in U.S. patent application Ser. No. 09/285,966 of R. Campbell et al, filed on Apr. 5, 1999 and assigned to the assignee of the present invention (and incorporated herein by reference). In that disclosure, a coolant stream moving through passage holes in a substrate is purposefully disrupted by the presence of an exit site on the "hot" side of the substrate. The exit site is preferably in the form of a crater, and may be contained within a thermal barrier coating applied to the hot side. (Passage holes are sometimes referred to herein as "film cooling holes").
One method of forming such a crater is described in U.S. Pat. No. 5,902,647 of M. Borom et al, which is incorporated herein by reference. In some embodiments of that invention, film cooling holes are temporarily filled with a masking material which is extruded into the holes from the backside (i.e., the "cold" side) of a turbine substrate. The masking material flows through the holes and exits at the hot side of the substrate. When the masking material exits the hole, it is cured, and forms a protrusion to which thermal barrier coatings do not adhere. After the coatings are applied, the masking material is removed to uncover the passage holes, which also results in the desired exit site geometry.
The processes described in the referenced patent applications are very suitable for increasing cooling effectiveness in many situations. However, new methods for enhancing the performance of the coolant stream are still desired in the art. The methods should be especially applicable to rows of film cooling holes in turbine engine components exposed to very high operating temperatures. More specifically, the methods should result in a film hole geometry in which cooling air exiting the holes adheres significantly to the hot surface of the substrate. These methods should also serve to minimize the undesirable mixing of cooling air with hot combustion gases.
Moreover, the new methods should not interfere with other functions of a particular component, e.g., the efficient operation of a turbine engine, or the strength and integrity of turbine engine parts. The methods should also be compatible with processes used to apply thermal barrier coating systems. In some instances, it would also be very desirable if the methods did not require access to the backside of a turbine component, e.g., an enclosed, internal section of the component. Finally, the implementation of these methods should preferably not involve a substantial cost increase in the manufacture or use of the relevant component, or of a system in which the component operates.